Space vehicle attitude control by microrockets utilizing subliming solid propellants



Oct. 6, 1970 M. E. MAEs 3,532,297

SPACE VEHICLE ATTITUDE CONTROL BY MICROROCKETS- UTILIZING SUBLIMING SOLID PROPELLANTS Original Filed Sept 24, 1963 4 Sheets-Sheet 1 m tt 2 INVENTOR. M/C'HEL MAE6 M. E. MAES Oct. .6, 1970 SPACE VEHICLE ATTITUDE CONTROL BY MICROROCKETS UT ILIZING NIS 4 Sheets-Sheet 2 SUBLIMING SOLID PROPELLA Original Filed Sept. 24, 1963 3,532,291 ILIZING M. E. MAES Oct. 6; 1910 SPACE VEHICLE ATTITUDE CONTROL BY MICROROCKETS UT SlBLIMING SOLID PROPELLANTS l9 3 4 Sheets-Sheet 5 Original Filed Sept. 24,

INVENTOR. M/C//L 6'.

Oct. 6, 1970 O M. E. MAES. 3,532,297

SPACE VEHICLE ATTITUDE-CONTROL BY MICRORQCKETS UTILIZING SUBLIMING SOLID PROPELLANTS Original Filed Sept. 24, 1963 4 Sheets-Sheet 4.

INVENTOR. M/Cl/EA MAL-5' United States Patent O US. Cl. 244-1 15 Claims ABSTRACT OF THE DISCLOSURE Satellite and like space vehicles incorporating jet reaction motor means arranged for attitude control, the motor means using a subliming solid as the propellant to obtain controlled thrust without chemical reaction or other molecular modification of the propellant, and with or without valvular control of propellant flow. A variety of suitable sublimable propellants are presented, which characteristically have a molecular Weight in vapor phase of less than about 108, a vapor pressure at 65 F. of at least about 0.5 p.s.i.a., a solid phase throughout a temperature range of from about 40 F. to about 120 F., a density in solid phase of at least about 0.05 lbs. mass per cubic inch, and a heat of vaporization of not more than about 1000 B.t.u./ lb. Typical examples are ammonium bicarbonate, ammonium carbamate, and ammonium hydrosulfide.

RELATED APPLICATION This application is a continuation of my now abandoned application Ser. No. 311,054, filed Sept. 24, 1963, and entitled Microrocket Systems Utilizing Subliming Solid Propellants. Application Ser. No. 459,685, filed May 28, 1965, and entitled Valveless Microrocket System, is a division of application Ser. No. 311,054.

FIELD TO WHICH INVENTION PERTAINS The present invention relates to space vehicle attitude control by means of small, non-combustion type reaction motors, and more particularly to in-orbit control of a space vehicle by low thrust rocket motor means mounted on the vehicle to effect changes in vehicular position in relation to the inertial path of travel of the vehicle, such motor means utilizing a self-feeding, subliming solid propellant, exhausted through reaction nozzle means, to produce reaction propulsion. Specific aspects of the invention pertain to various of the factors involved as to selection of the propellant for such so-called subliming solid motors, as to design of system components, 'and as to performance characteristics of rocket motors of this type.

NATURE OF PRIOR ART Mission studies and control system analyses have shown that only very little thrust is required to control vehicular attitude, orbital position, or spin rate of satellites or like space vehicles. In general, such vehicular control of satellites and the like is best performed with a maximum of propellant economy through use of what may be termed microrockets, i.e. a rocket producing thrust on the order of about to 10- pounds thrust.

Prior to the present invention, the only practical type of operationally available propellant system for satellite attitude control utilizes cold nitrogen gas as the propellant. The major disadvantages of the cold nitrogen gas system are that it requires a relatively high system weight due to the low density of the propellant and requires a 3,532,297 Patented Oct. 6, 1970 high pressure in the propellant storage tank (typically about 3,000 p.s.i.a. initially), coupled with the additional weight of the necessary pressure regulator means and control valves. characteristically, an attitude control system involving a cold nitrogen gas jet reaction motor requires that the weight of the propellant storage tank at least equal the initial weight of the propellant. Another major disadvantage of the 'cold gas reaction motor for this purpose is the fact that the system contains a multiplicity of moving parts, operating under high pressure, often giving rise to considerable leakage and a low reliability with regard to satisfactory operation over extended periods.

SUM-MARY OF THE INVENTION FEATURES It is a primary feature and advantage of subliming solid microrocket motors for vehicular attitude control according to the present invention that such motors offer better performance, higher reliability, and markedly lower equipment weight than comparable cold nitrogen gas reaction jet motors, the subliming solid rocket motor usually requiring a propellant tank weight which is only onetenth or less of the initial weight of the propellant. Further, the subliming solid microrocket is compact, involves simple components, is capable of providing higher specific impulse than a cold nitrogen gas system, has a greater propellant density, a generally much lower operating pressure (the exhaust pressure of the subliming solid system in most cases being 15 p.s.i.a. or lower), is stable under storage, and has a longer operating life capability. In addition, in contrast to combustion type bi-propellant reaction jet motors, the subliming solid microrocket motors as employed for attitude control according to the present invention obviate any need for either propellant ignition or propellant combustion. Further, subliming solid microrocket motors as employed for satellite attitude control according to the present invention are self-feeding and are particularly suited to the intermittent pulsing mode of operation necessary for satellite attitude control.

These and other objects, features, advantages and characteristics of the invention will be apparent from the following discussion of various suitable subliming solid propellants, component designs, and motor characteristics, taken together with the accompanying drawings illustrating certain typical and therefore non-limitive embodiments of satellites incorporating such motors.

DESCRIPTION OF THE DRAWINGS In the drawings, wherein like letters and numerals refer to like parts:

FIG. 1 is an isometric view, with various portions broken away to show interior detail, of a satellite, incorporating a subliming solid type control rocket means with four orthogonally related nozzles, control of propellant delivery to each of the nozzles being effected by separately actuatable valves;

FIG. 2 is an isometric, cutaway view similar to the View of FIG. 1 showing a satellite with modified form of subliming solid type microrocket motor means, incorporating an auxiliary heating element within the propellant tank;

FIG. 3 is a further isometric cutaway view similar to FIGS. 1 and 2, showing a satellite with a further modified form of subliming solid type microrocket motor means, incorporating two oppositely directed nozzles and a thermally controlled variable orifice functioning to maintain a substantially constant propellant mass flow during variation in propellant temperature;

FIG. 4 is an enlarged detail view partially in cross-section and partially in elevation, showing further solenoid valve and nozzle detail of the nozzle assembly incorporating in the subliming solid type microrocket motor means shown in FIG. 3, and further illustrating a suitable technique for maintaining the valve and nozzle at somewhat higher temperature than the propellant to inhibit propellant recondensation;

FIG. 5 is an isometric view, with portions broken away to show interior detail, of a further modified form of subliming solid type microrocket motor means according to the present invention, specially adapted for spin control of the satellite, with oppositely related nozzles fed from a common bidirectional control valve, the system shown in FIG. 5 also incorporating a modified form of flow path heating means;

FIG. 6 is a fragmentary view of one of the propellant conduits of the rocket motor means shown in FIG. 5, illustrating a further modified form of flow path heating means therefor;

FIG. 7 is an isometric cutaway view of yet another form of subliming solid type microrocket motor means characterizing the invention, with valveless control of propellant flow by means of selective radiant heating of the propellant surface, and

FIG. 8 is an isometric cutaway view of another form of valveless subliming solid type microrocket motor means wherein propellant fiow through the nozzle is controlled by heating of the propellant surface through the admission of solar energy, providing a completely passive control system.

DETAILED DESCRIPTION The form of subliming solid type microrocket motor means shown in FIG. 1 comprises a propellant tank 10 of spherical configuration and constructed of a suitable lightweight material such as aluminum or glass fiber reinforced resin, for example. The subliming solid propellant, in the microrocket shown at FIG. 1, is a compressed powder cake 12 which, as will be readily understood, initially occupies substantially all of the tank 10 and gradually reduces in size by surface sublimation, the propellant cake 12 shown in FIG. 1 being in a partially depleted state. Mounted at the top of the propellant tank .10 is a thrustor block 14 in which are threadedly mounted four valve and nozzle assemblies comprising respective solenoid valves 16, 18, 20, 22, each in communication with respective nozzles 24, 26, 28, 30. Each of the valves 16, 18, 20, 22 is a solenoid actuated coaxial type valve, conventional per se, comprising (as shown in the cutaway view of valve 16) a spring-loaded plunger 32 actuated away from the valve seat 34 by energization of solenoid coil 36. Each of the nozzles 24, 26, 28, is per se of the conventional conical type. As will be apparent in FIG. 1,

the flow of propellant vapor from the tank 10 into the thrustor block 14 is through threaded fitting 38 into the interior chamber 40 of the thrustor block 14, thence out Whichever valve and nozzle assembly is open. In order to prevent any solid particle movement from the propellant tank 10 into the valve and nozzle assemblies, a filter screen 42 is seated in the fitting 38. Fitting 38 on which the thrustor block 14 is mounted is in turn attached to the tank 10 by threaded engagement with outlet port 44 Welded to the tank 10, and a convenient manner of mounting the microrocket motor tangentially of the satellite which it controls is to clamp the skin or casing of the satellite between the tank outlet port 44 and the fitting 38, a fragment of the satellite casing being shown so assembled, at S.

In order to be compatible with the propellant vapor, the various flow path means for the vapor are constructed of a material not subject to chemical attack by the vapor, such as aluminum or stainless steel.

FIG. 2 illustrates a modified form of the subliming solid control rocket shown in FIG. 1, incorporating all of the components discussed above and further including a selectively energized electric heating element 46 threadedly mounted in the propellant tank 10 and extending into the tank 10 so as to radiantly heat the propellant cake 12 in the event operating conditions are such that auxiliary heating of the propellant is desirable or required to main- Cit tain operating vapor pressure. One of the operating characteristics of the subliming solid microrocket motor is that heat is required to cause the propellant to sublime, i.e. the propellant has a certain heat of sublimation requirement. In general, the necessary heat of sublimation is attained from the ambient environment by thermal conduction and radiation through the propellant tank. However, in those cases where adequate heat cannot be drawn from the surroundings to maintain the propellant at desired operating temperature, the auxiliary heating element 46 prevents the temperature of the propellant cake 12 from falling below the desired operating level. Heating element 46 thus serves to in efifect extend the maximum time during which the microrocket motor can be operated continuously.

FIG. 3 illustrates further typical variations in components of subliming solid type microrocket motors according to the invention. In the form of rocket motor shown in FIG. 3, the generally spherical propellant tank 10A is partially filled with subliming solid propellant 12A, in this case in loose powder form, and the control valve nozzle assemblage comprises thrustor block 50 and respective oppositely related valves 52, 54 and nozzles 56, 58. The said valves 52, 54 are of the shear seal type, conventional per se, selectively controlled by respective solenoids 60, 62. The specific construction thereof is shown in FIG. 4 and discussed in more detail below. Nozzles 56, 58 are of a conical type, conventional per se. Flow of propellant vapor from the tank 10A to the thrustor block 50 is through a filter screen 64 and through a thermally controlled variable orifice assembly 66, then through mounting tube 68 into the manifold chamber 70 (see FIG. 4) of the thrustor block 50. The thermally controlled variable orifice assembly 66 is threadedly mounted to the top of tank 10A and the functional portions of this assembly comprise a pintle 72 facing the contoured internal wall 74 of the orifice chamber surrounding the pintle 72. Said pintle 72 is linked by a flexible push rod 76 in a flexible sheath 78 leading to a temperature sensing expansion element 80 within the solid propellant 12A. The temperature sensing expansion element 80 and the thermally controlled variable orifice assembly 66 constitute a commercially available unit, conventional per se, and the operation thereof is such that with an increase in temperature at the sensing element 80, element 80 moves push rod 76 which in turn moves pintle 72 to relatively reduce the orifice area between the pintle 72 and the contoured wall 74, with the result that on the occasion of propellant flow the flow rate of propellant vapor through the orifice and into the thrustor block 50 is maintained substantially constant in spite of the change in vapor pressure of the propellant caused by change in temperature.

As an operational refinement of microrocket motors according to the invention, such as that shown at FIG. 3, nitrogen gas can be initially placed in the space 82 in tank 10A not occupied by solid propellant 12A, at moderate pressure (about 150 p.s.i.a., for example), in order to extend the initial pulse duration of the system without regard to providing heat of sublimation, such as often required during initial acquisition of desired satellite attitude immediately after reaching orbital velocity. In this manner, a single attitude control system can be made to optimally perform both the functions of initial attitude acquisition and of long term attitude control.

FIG. 4 presents an enlarged detail view of a portion of the thrustor block 50 and the valve 52, nozzle 56 assembly shown in FIG. 3. As shown, the valve 52 is of the shear seal type, opened upon energization of solenoid 60, movement of the shear seal block 84 upwardly (as viewed) permitting propellant vapor flow from chamber 70 of the thrustor block 50 through thrustor block passage 86 and valve passages 88, 90 into the reaction nozzle 56.

One of the operational considerations with respect to a subliming solid type rocket motor is that all How paths for the propellant vapor are in many cases to be maintained at a temperature which inhibits recondensation of the propellant vapor. One simple way to do this is to maintain the flow paths at a slightly higher temperature than the recondensation temperature of the propellant vapor, and as illustrated in FIG. 4 with respect to the block 50, valve 52 and nozzle 56, for example, is to provide the flow path exterior surfaces, or at least portions thereof, with a surface coating having a coefficient of thermal absorptivity causing preferential heating of the fiow path elements. For this purpose, the valve and nozzle assembly shown in FIG. 4 includes a thermally absorptive coating, typically a flat black paint, as indicated at 92.

Anotherway in which the propellant vapor flow path means can be conveniently heated, particularly under conditions where the flow path means are exposed to a low temperature environment compared to the temperature within the propellant tank, is to provide the necessary heat by means of a small continuous current passing through the solenoid coils of the flow path valve means at a level below that required for solenoid actuation. A power input on the order of a few tenths of a watt, applied in the manner, is sufficient to maintain the valve and associated components at appropriate temperature even when the flow path components are exposed to the black background of deep space. Thus, a small heating current through the coils of solenoids 60, 62 in the system illustrated at FIG. 3, or through the solenoids of valves 16, 18, 20, 22 in the systems illustrated at FIGS. 1 and '2, can also serve to provide the flow path temperature condition desired to inhibit propellant recondensation. As will be apparent, this technique can be applied in conjunction with other forms of fiow path heating means, such as the temperature absorptive coating illustrated in FIG. 4, for example.

FIG. 5 is an isometric view, with certain portions broken away to show interior detail, of another typical component arrangement for subliming solid type microrocket motors according to the invention, specifically configured to provide pure-couple spin control. In this system, the propellant tank, generally indicated at B, comprises cylinder with a closed bottom and a dished cover 102 bolted in place thereon to provide a pressure enclosure for a packed powder cake 12B of subliming solid propellant. Exhaust of propellant vapor from the propellant tank 10B is controlled by means of a three-position shear seal valve 104, controlled bidirectionally by an electrically energized torque motor 106. The said valve 104 having three control positions, delivering propellant vapor to either outlet conduit 108 or outlet conduit 110 depending upon the polarity of energization of the torque motor 106, with a center control position preventing propellant flow to either outlet conduit when no energization of the torque motor 106 is provided. The three-way valve 104- and torque motor 106 shown in FIG. 5 is a commercially available unit, conventional per se, and accordingly need not be further described for an understanding of the operation thereof.

The fiow path arrangement for propellant vapor delivered to conduit 108 comprises a T-fitting 112 and conduits 114, 116, respectively in communication with oppositely directed conical nozzles 118, through the passageways provided in respective thrustor blocks 122, 124, each identical with the other, the nozzle 120 and its associated passageway in thrustor block 124 being cut away in FIG. 5 to show some of the interior detail thereof. Similarly, propellant vapor delivered to conduit 110 flows through T-fitting 126 and respective conduits 128, 130 to the thrustor blocks 122, 124 and nozzles 132, 134.

Mounting of the microrocket motor shown in FIG. 5 tangentially of the satellite which it controls can be by suitable attachment of the thrustor blocks 122, 124 externally of the satellite casing, indicated fragmentarily at SA, the microrocket motor components internally of the satellite being supported by suitable structure means, not shown.

The microrocket motor shown in FIG. 5 is further characterized by utilization of a different form of propellant vapor flow path heating means than that shown in FIG. 4. Specifically, by way of further example, with respect to such flow path heating means, the various conduits 108, 110, 114, 116, 128, 130 are each provided with a sleeve 136 or the like containing radioactive material such as Pm or Pu Fission of the radioactive material provides a degree of thermal heating of the various conduits and associated components in thermally conductive association therewith. As will be apparent, the radioactive material to accomplish this purpose can be present in various forms other than the sleeves 136 shown, such as by being applied as internal or external films or coatings in or on the conduits or portions thereof, or as by being placed within suitable compartmentation in the valve 104 and/or the thrustor blocks 122, 124, simply by way of further example.

FIG. 6 is a fragmentary isometric view showing a further form of flow path heater means which can be used in lieu of the radioactive sleeves 136. In this heating arrangement an electric sleeve heater, suitably energized, thermally heats each conduit, such as shown in FIG. 6 at 138 with respect to conduit 116.

FIG. 7 typically illustrates what may be termed a subliming solid type microrocket motor of the valveless type. In this form of rocket motor, the control of propellant vapor flow is solely by means of selective heating of the propellant. More specifically, in the rocket shown in FIG. 7, a generally cylindrical, round-ended propellant tank 10C is partially filled with a compacted cake 12C of subliming solid propellant, with an insulating sheath 140 being provided between the tank and propellant to prevent heat transfer to the propellant cake 120 through the wall of tank 10C. A conical reaction nozzle 142 is attached to the tank 10C at a point thereof above the propellant cake 12C, the said nozzle 142 being in direct and open communication with the chamber 144 above the propellant cake 12C. Selective heating of the exposed surface of propellant cake 12C is by means of electric heating element 146, suitably energized on command, as by means of solid state power switch 148, of a type conventional per se, applying a power input from conductors to the heating element 146 through conductors 152 at such time as the normally open power switch 148 receives energization from a control input applied to conductors 154. A generally hemispherical reflector 156 is preferably employed, to direct the radiant energy from heating element 146 onto the surface of the propellant cake 12C. In a case where the propellant cake 12C is relatively light in color, a small amount of finely divided particles which are darker in color, such as carbon black, can be dispersed in the propellant cake 12C in order to improve the heat absorption characteristics of the propellant cake surface.

The mode of operation of a valveless form of subliming solid type microrocket motor such as shown at FIG. 7 is as follows. Assuming that the propellant cake 120 is at an initially ambient temperature, an initial vaporization of propellant occurs until by the loss of heat of sublimation the thermally insulated propellant cake is refrigerated to the point where the vapor pressure of the propellant is so low that substantially no propellant flow exists. This condition continues to exist as long as there is no heat input to the propellant cake. Since there is substantially no propellant flow and no thrust is being produced, this condition can be referred to as a nonoperative or off condition. At such time as an operating pulse is desired, the heating element 146 is energized and the heat input to the surface of the self-refrigerated propellant cake 12C causes localized heating of the surface thereof, producing propellant vapor by surface sublimation and vapor flow out of the open nozzle 142. This thrust producing flow of propellant vapor out the nozzle 142 can be referred to as the operating or on condition, which condition will continue substantially only as long as the heating element 146 remains energized. Thus, the valveless microrocket shown in FIG. 7 can be pulsed for any desired period simply by selective energization of the heating element 146. A primary feature and advantage of the valveless microrocket motor is that pulses of thrust are produced on command without the use of moving parts to control propellant flow. The absence of any moving parts in the system has the capability of providing extremely high unit reliability. It will be apparent that, in a valveless microrocket motor such as shown in FIG. 7, any type of means for selectively producing energy other than electrical heating element 146 can be used to selectively generate propellant flow. By way of further examples, the selectively controllable radiant energy producing means can be; (a) a chemical reaction type heat source, such as a small oxyhydrocarbon burner, preferably with the heat produced products of combustion being separately exhausted without introduction to the propellant tank; (b) a dielectric heating means, such as spaced dielectric plates arranged within the solid propellant or at spaced points on the insulative wall 140; (c) electroconductive heating means, such as electrodes immersed in and passing electrical current directly through the solid propellant mass; or (d) a laser type device, such as silicon carbide crystal laser controlled by the quantity of electrical current flow through the crystal.

Yet another technique for selectively heating the surface of a valveless subliming solid type microrocket motor to produce attitude controlling thrust on demand is to heat the surface with solar energy. An example of a valveless microrocket motor embodying the principle 3 is presented at FIG. 8. In FIG. 8, the tank C, propellant cake 12C, insulative liner 140 and nozzle arrangement 142 are arranged in the same manner as in the valveless rocket motor illustrated at FIG, 7, and correspondingly so designated. To provide selective heating of the exposed surface of propellant cake 12C, the end of tank 10C facing the propellant surface is provided with a light transmitting window such as quartz lens 158. As an application of this type of microrocket motor for satellite attitude control, the motor can be oriented on the satellite in such a manner that, when correct satellite attitude is obtained, no incident sunlight strikes the quartz lens 158 and no thrust is produced by the motor. However, should the satetllite move out of correct attitude, the quartz lens 158 receives incident sunlight and the solar energy transmitted to the quartz lens onto the surface of the propellant cake 12C causes surface sublimation of the propellant and the motor produces thrust, returning the satellite to correct attitude, at which time the sunlight no longer strikes the quartz lens and the thrust is interrupted. It is to be noted that in this form of valveless microrocket motor, the control of thrust generation is entirely passive, i.e. requires no control energy producing means and no active control elements on board the satellite.

As will 'be evident, a suitable shield and optical system can be used in conjunction with the light transmitting quartz lens 158 to bring incident solar energy onto the solid propellant only when attitude correction is desired. For example, angle of incidence of solar energy operable to produce attitude corrective thrust can be narrowed and can have sharp angular demarkation, as by use of a light shielding tube (not shown) around the lens 158, with a slotted or like aperture at the tube end opposite from the lens.

While the term valveless has been used to describe the type of subliming solid type rocket motor such as shown at FIGS. 7 and 8 where thrust generation is controlled solely by control of the energy input to the subliming solid propellant, it will be understood that use of a simple spring-loaded over pressure relief valve or the like in the propellant vapor flow path between the chamber 144 and the reaction nozzle 142, to permit exhaust of propellant vapor only when the vapor pressure is above a certain level (eg 0.5 p.s.i.a.), is not inconsistent with the essential valveless mode of propulsion control characteristic of this type of rocket motor and is to be considered within the scope of the present invention.

Having considered the component makeup of various subliming solid type jet reaction motors according to the invention, consideration will next be given to the matter of subliming solid propellant selection, including an indication of the basic properties of suitable propellants and the chemical constituency thereof.

As to the criteria for propellant selection, the important properties to be considered are vapor pressure, melting point, molecular weight of the propellant vapor, density, heat of sublimation, specific impulse, and chemical stability. With regard to vapor pressure, the subliming solid propellant must possess a significant vapor pressure at operating temperature and the operating vapor pres sure should fall within certain limits to obtain optimum performance, which limits depend to a considerable extent upon the thrust level under consideration. For example, if the vapor pressure of a propellant is too high, relative to the thrust level, the throat area of the exhaust nozzle is unduly small and might tend to be obstructed should operating conditions vary considerably from those for which the system was designed. If the vapor pressure is too low, on the other hand, the nozzle throat area becomes excessively large, leading to undue weight of the nozzles, flow path conduits and valves. By way of practical limits with respect to nozzle throat diameters and vapor pressures, the propellant vapor pressure should be at least about 0.5 p.s.i.a. and preferably not over about 15 p.s.i.a.

With respect to the melting point of the subliming solid propellant, the melting or decomposition point of the propellant cannot be too low, otherwise the propellant will melt under unexpected environmental conditions resulting in higher than design temperatures. The propellant should be specifically selected for each application to insure compatibility between the propellant melting point and the expected satellite environmental condition. In general, the melting point is to be at least about 120 F.

With respect to molecular weight of the propellant vapor, it is necessary to have a reasonably good specific impulse (I i.e. at least about 40 and preferably at least about lb. f./lb. m./sec., in order to obtain good performance from a subliming solid propellant system. As is known, the primary physical quantity which affects the specific impulse (pounds of thrust produced per unit weight fiow of propellant per second) is the molecular weight of the exhaust gases. It is desirable to have the molecular weight of the exhaust gases as low as possible. In this respect it is to be emphasized that most substances which have an appropriate vapor pressure are impractical because of the fact that the molecular weight of the substance in vapor phase is undesirably high; this is often particularly true in the case of volatile organic materials such as napthalene. It is thus an important characteristic of most subliming solid propellants suitable for purposes of the present invention, as listed in Table I below, that they reversibly decompose upon sublimation from the solid phase to form a gas phase mixture of lighter compounds. Thus, for example, ammonium hydrosulfide reversibly becomes ammonia and hydrogen sulfide in gas phase, and ammonium bicarbonate reversibly becomes ammonia water vapor, and CO in gas phase, with these respective gas phases having molecular weights of 25.6 and 26.4, for example. In general the molecular weight of the subliming propellant vapor should be not more than about 108, preferably should 9 be less than about 60, and optimally should be less than about 30.

As to the density of the subliming solid propellant, it is desirable that it be as dense as possible in order to reduce the weight of the propellant tank for a given amount of propellant. In general, the density of the subliming solid propellant should be at least about 0.05 pound mass per cubic inch.

The heat of vaporization of the propellant should be as low as possible, since this directly determines the allowable pulse length and maximum duty cycle as a function of the thrust level and amount of propellant in the tank. As will be evident, if the heat of vaporization is excessively large, then the self-cooling effect during the pulse becomes more pronounced. In general, the heat of vaporization of the subliming solid propellant should not exceed about 1000 B.t.u./lb.

It is also important that the propellant material exhibit chemical stability under operating and storage conditions and should not have corrosive, toxic or other undesirable characteristics unless such can be tolerated in the environment of use.

The following is a tabulation of certain suitable subliming solid propellants for use in microrocket motors according to the present invention, together with an indication of the various above-discussed physical and chemical properties of the propellants.

From the foregoing considerations, it will be evident that other variations, modifications and adaptations of subliming solid propellant type jet reaction motors can be made. Thus, by way of further typical example, the subliming solid propellant in the form occupying the propellant tank 10 can in certain cases be crystalline or amorphous rather than particulate or powdered, and can be configured to present other than a flat exposed surface. Also, consistent with the controlling physical and chemical properties, it will be understood that other solid subliming propellants are possible, besides those listed in Table I. As to those microrocket motors employing valve type control for propellant vapor flow, other forms of valves can be employed, other than the specific solenoid actuated coaxial and shear type valves disclosed. As will also be apparent, any desired reaction nozzle configuration can be employed other than the conical form of nozzle disclosed, such as the well-known bell type reaction nozzle. As to the manner of heating the solid propellant in the propellant tank, it will also be understood that any other form of suitable energy source can be employed, such as the energy prOducing means above discussed in connection with the form of rocket shown at FIG. 7. Yet another specific example in this respect, applicable Where a continuous (i.e. non-controlled) heating of the subliming solid propellant is desired to main- IABLE l.-PROPERTIES OF SUBLINTNG SOLID PROIELLANTS Vapor pressure Melting Molecular Heat of Approx. Theo. Chemical at 65 F, Point weight of Density sublimation specific irnstability subliming Solid Propellaut Formula (lb. i. in. F.) gas phase (lb. m./1n (l3.t.u. .lb.) pulse (at 65 F.)

Ammonium bicarbonate NI'IJIICOS 0. G2 150 26. 4 0. 0572 920 84 Stable. Ammonium carbonate (NIIOEOOK 0. 6 150 0 .0a 3b Ammonium earbainate NI'IJCOQNI'IZ 1. 16 150 0 84 Do, Ammonium hydrosulfirle NHJIS 6.37 150 25. 6 85 Do. Ammonium sulfide (NHmS 7.0 150 22.8 37 1) Cyanogen bromide CNBr 1. 55 124 105. 0 42 Do. Phosphonium bromide IILBr 4. 2 150 57. 57 DO, lliosphouium io(lide 1. 1 143 81. 0 47 Do. Sulfur trioxide 0. 87 1M 80- 0 48 Do.

Since sublimation of each of the propellants listed in Table I is essentially a physical phenomenon, it will be understood that any chemically non-reactive mixtures of subliming solid propellants can also be employed if desired.

The following specific examples typify some of the design considerations involved in designing subliming solid microrocket motors for particular applications at various thrust levels. For a subliming solid microrocket to operate at a thrust level of l0 lb. at a temperature of 65 F., one good subliming solid propellant is ammonium carbonate, having a vapor pressure of 0.6 lb. f./in. at the assigned operating temperature. At this operating pressure and thrust level, a suitable nozzle throat diameter is 0.002 inch and a suitable nozzle area ratio is 100. A microrocket motor designed according to these specifications has a continuous duty cycle capability.

For a subliming solid type microrocket motor specifically designed for operation as a valveless rocket such as shown in FIG. 7 or FIG. 8, with a thrust level of 10- 1b., a suitable subliming solid propellant is ammonium bicarbonate, providing an operating pressure of 10 p.s.i.a. with operating temperatures of 50 F. (off condition) and +l F. (on condition). In this system, a suitable nozzle throat diameter is 0.003 inch, with a nozzle area ratio of 100. A microrocket so designed has a continuous duty cycle capability.

For a subliming solid microrocket to operate at a thrust level of 10 1b., a suitable propellant is ammonium hydrosulfide, providing operating pressure of 6.37 p.s.i.a. at 65 F., a suitable nozzle throat diameter in this case being 0.033 inch, with a nozzle area ratio of 100. The duty cycle capability of this microrocket motor is 10%.

tain a relatively high level propellant vapor pressure, is the utilization of a radioactive material containing loose pellet or the like placed in the propellant tank along with a propellant.

What is claimed is:

1. In a space vehicle incorporating attitude contro means comprising jet reaction motor means having reaction nozzle means arranged to exert thrust on the vehicle to change the position thereof in relation to the inertial path of travel of the vehicle; the improvement wherein said jet reaction motor means comprises:

(a) a sublimable solid propellant having a solid state density of at least about 0.05 pound mass per cubic inch and a solid state throughout a temperature range of from about 40 F. to about F., a molecular weight in vapor phase of not more than about 108, and a substantial vapor pressure at 65 R, such sublimable propellant being characterized by loss of heat of sublimation incident to vaporization thereof so as to be self-cooling to the point where the vapor pressure of the propellant is so low that substantially no propellant vaporizes in the absence of an external heat input;

(b) a propellant container confining said propellant and including means providing an external heat input to said propellant;

(c) flow path means communicating said container with said reaction nozzle means;

(d) solid particle barrier means in said flow path; and

(e) valve means in said flow path means, downstream of said solid particle barrier means, for controlling the extent of flow of vaporized propellant from said container to said reaction nozzle means.

2. The space vehicle according to claim 1, wherein the 1 1 means providing an external heat input to said propellant comprises radiant heater means spaced from and radiantly heating said propellant.

3. The space vehicle according to claim 2, wherein said heater means comprises a selectively energizeable electric heating element.

4. The space vehicle according to claim 1, said propellant being selected from the group consisting of ammonium bicarbonate, ammonium carbonate, ammonium carbamate, ammonium hydrosulfide, ammonium sulfide, cyanogen bromide, phosphonium bromide, phosphonium iodide, sulphur trioxide, and chemically non-reactive mixtures thereof.

5. The space vehicle according to claim 4, wherein said propellant principally comprises ammonium bicarbonate.

6. The space vehicle according to claim 4, wherein said propellant principally comprises ammonium hydrosulfide.

7. The space vehicle according to claim 1, further comprising moderately pressurized gas in the propellant container along with the subliming solid propellant and functioning to enhance the initial thrust of the motor means without regard to providing heat of sublimation for the solid propellant.

8. The space vehicle according to claim 1, wherein said reaction nozzle means is directed generally tangentially of the vehicle.

9. The space vehicle according to claim 8, wherein the means providing an external heat input to said propellant comprises radiant heater means spaced from and radiantly heating said propellant.

10. The space vehicle according to claim 9, said propellant being selected from the group consisting of am monium bicarbonate, ammonium carbonate, ammonium carbamate, ammonium hydrosulfide, ammonium sulfide, cyanogen bromide, phosphonium bromide, phosphonium iodide, sulphur trioxide, and chemically non-reactive mixtures thereof.

11. The space vehicle according to claim 10, further comprising moderately pressurized gas in the propellant container along with the subliming solid propellant and functioning to enhance the initial thrust of the motor means without regard to providing heat of sublimation for the solid propellant.

12. The space vehicle according to claim 8, said propellant being selected from the group consisting of ammonium bicarbonate, ammonium carbonate, ammonium carbamate, ammonium hydrosulfide, ammonium sulfide, cyanogen bromide, phosphonium bromide, phosphonium iodide, sulphur trioxide, and chemically non-reactive mixtures thereof.

13. The space vehicle according to claim 8, further comprising moderately pressurized gas in the propellant container along with the subliming solid propellant and functioning to enhance the initial thrust of the motor means without regard to providing heat of sublimation for the solid propellant.

14. The space vehicle according to claim 13, said propellant being selected from the group consisting of ammonium bicarbonate, ammonium carbonate, ammonium carbamate, ammonium hydrosulfide, ammonium sulfide, cyanogen bromide, phosphonium bromide, phosphonium iodide, sulphur trioxide, and chemically non-reactive mixtures thereof.

15. A low thrust rocket system comprising a propellant container, reaction nozzle means, flow path means communicating said container with said reaction nozzle means, and a sublimable solid propellant in said container, said propellant having a solid state throughout a temperature range of from about 40 F. to about 120 F, a molecular weight in vapor phase of not more than about 108, and a vapor pressure at 65 F. of at least about 0.5 p.s.i.a., said rocket system further comprising moderately pressurized nitrogen gas or the like contained in the propellant container along with the subliming solid propellant and functioning to extend the initial pulse duration of the system without regard to initially providing heat of sublimation for the solid propellant.

References Cited UNITED STATES PATENTS 2,671,312 3/1954 Roy 60--39.48 2,816,419 12/1957 Mueller 6037 3,024,941 3/ 1962 Vandenberg. 3,097,480 7/1963 Sohn. 3,097,818 7/1963 Heller. 3,159,967 12/1964 Webb 60-202 FERGUS S. MIDDLETON, Primary Examiner US. Cl. X.R. 60200, 229 

0.5 P.S.I.A., A SOLID PHASE THROUGHOUT A TEMPERATURE RANGE OF FROM ABOUT 40* F. TO ABOUT 120* F., A DENSITY IN SOLID PHASE OF AT LEAST ABOUT 0.05 LBS. MASS PER CUBIC INCH, AND A HEAT OF VAPORIZATION OF NOT MORE THAN ABOUT 1000 B.T.U./ LB. TYPICAL EXAMPLES ARE AMMONIUM BICARBONATE, AMMONIUM CARBAMATE, AND AMMONIUM HYDROSULFIDE. 